Compositions and methods of producing rocket propellants with enhanced cryogenic cooling, thermal stability, and thrust efficiency performance

ABSTRACT

The present disclosure relates to a rocket propellant composition comprising (a) a liquid oxygen; and (b) a hydrocarbon mixture comprising: (i) a hydrogen content from about 14.0 mass % to about 16.0 mass %, by mass of the hydrocarbon mixture; (ii) a kinematic viscosity of from about 5 mm 2 /s to about 8 mm 2 /s at a temperature ranging from about −35° C. to about −33° C.; and (iii) a sulfur content of from about 0 ppm to about 0.1 ppm, by mass of 0 the hydrocarbon mixture.

FIELD OF THE DISCLOSURE

The present disclosure relates, in some embodiments, to rocket propellants having improved fuel thrust and cryogenic cooling metrics and methods of making same.

BACKGROUND OF THE DISCLOSURE

Rocket propellants can be categorized as either solid propellants or liquid propellants. While solid propellants may be easier to store, they have a significant disadvantage of having a lower specific impulse, which is a measure of propellant efficiency, resulting in an overall lower performance in comparison to their liquid counterparts. Besides having a higher specific impulse, liquid propellants are capable of being throttled in real time, shut down, and restarted, whereas once a solid propellant is ignited, it must burn until all of the propellant is used.

Currently used liquid propellants include Rocket Propellant-1 (RP-1) and Rocket Propellant-2 (RP-2)(RP-2 being a lower sulfur version of RP-1), which are highly reformed forms of kerosene that have both a low specific impulse and is stable at room temperature. Generally, both RP-1 and RP-2 are combusted in a rocket engine with liquid oxygen as the oxidizer. However, given that the weight of a fuel carries with it a heavy burden on rocket propulsion efficiency, new advancements need to be made to discovery denser and more efficient fuels such as those having a higher hydrogen content than existing liquid propellants. Additionally, given that liquid propellants are generally cooled before being combined with liquid oxygen, advanced liquid propellants are needed that can be cooled to low temperatures while having lower viscosities than existing liquid propellants so they can be properly combined with liquid oxygen when needed. Also, since existing liquid propellants tend to produce coke while being consumed, which may lead to engine fouling, liquid propellants are needed having a reduced potential to form coke and working temperatures.

SUMMARY

Accordingly, there is a need for improved liquid rocket propellant compositions having a high hydrogen content and a low kinematic viscosities at temperatures below −33° C. In some embodiments, the present disclosure relates to a rocket propellant composition including a liquid oxygen and a hydrocarbon mixture. During a combustion process, a liquid oxygen and a hydrocarbon mixture are combined and then combusted to create propulsion for a rocket.

The higher a hydrogen content value is of a hydrocarbon mixture, the higher of a specific impulse the hydrocarbon mixture can generally provide for a rocket, which is a measurement of how effectively the rocket can use the hydrocarbon mixture. A disclosed hydrocarbon mixture has a hydrogen content according to ASTM D5291 ranging from about 14.0 mass % to about 16.0 mass %, preferably from about 15.0 to about 16.0 mass % by mass of the hydrocarbon mixture. A disclosed hydrocarbon has a relatively low kinematic viscosity according to ASTM D7042 and/or ASTM D445 of from about 5 mm²/s to about 8 mm²/s, preferably from about 5 to about 7 mm²/s at a temperature ranging from about −35° C. to about −33° C. Additionally, a hydrocarbon mixture includes a sulfur content according to ASTM D5623 of from about 0 01 ppm to about 0.1 ppm, by mass of the hydrocarbon mixture. Each of these attributes work synergistically to provide for better fuel efficiency while being used by a rocket in a combustion process.

In some embodiments, the present disclosure relates to methods of producing a rocket propellant from a liquid oxygen and a hydrocarbon mixture. A method includes providing a liquid oxygen and providing a hydrocarbon mixture. A hydrocarbon mixture includes a hydrogen content according to ASTM D5291 ranging from about 14.0 mass % to about 16.0 mass %, preferably from about 15.0 to about 16.0 mass % by mass of the hydrocarbon mixture, a kinematic viscosity according to ASTM D7042 and/or ASTM D445 of from about 5 mm²/s to about 8 mm²/s, preferably from about 5 to about 7 mm²/s at a temperature ranging from about −35° C. to about −33° C., and a sulfur content according to ASTM D5623 of from about 0.01 ppm to about 0.1 mass %, by mass of the hydrocarbon mixture. A liquid oxygen and a hydrocarbon mixture may be combined and then combusted.

BRIEF DESCRIPTION OF THE DRAWINGS

Some embodiments of the disclosure may be understood by referring, in part, to the present disclosure and the accompanying drawings, wherein:

FIG. 1 is a plot of the percent change in useful payload versus the percent change in specific impulse, according to a specific example embodiments of the disclosure;

FIG. 2 is a plot of specific impulse gain versus the liquid oxygen:hydrocarbon mixture ratio at 350 psi, according to a specific example embodiment of the disclosure;

FIG. 3 is a plot of specific impulse gain versus the liquid oxygen:hydrocarbon mixture ratio at 500 psi, according to a specific example embodiment of the disclosure;

FIG. 4 is a plot of specific impulse gain versus the liquid oxygen:hydrocarbon mixture ratio at 750 psi, according to a specific example embodiment of the disclosure; and

FIG. 5 is a plot of kinematic viscosity versus temperature of various disclosed rocket propellant compositions, according to a specific example embodiments of the disclosure.

DETAILED DESCRIPTION

The present disclosure relates, in some embodiments, to a rocket propellant composition and methods of making the rocket propellant composition. A disclosed rocket propellant composition has advantageous cryogenic cooling, thermal stability, and thrust efficiency performance with respect to existing rocket propellants. Applicants' rocket propellant composition achieves this by varying formulation components of a hydrocarbon mixture including a hydrogen content, a kinematic viscosity, and a sulfur content.

Rocket Propellant Compositions

According to some embodiments, a rocket propellant composition includes a liquid oxygen and a hydrocarbon mixture having advantageous cryogenic cooling, thermal stability, and thrust efficiency performance with respect to existing rocket propellants. A disclosed hydrocarbon mixture includes a hydrogen content according to ASTM D5291 ranging from about 14.0 mass % to about 16.0 mass %, preferably from about 15.0 to about 16.0 mass % and a sulfur content according to ASTM D5623 of from about 0.01 ppm to about 0.1 ppm, by mass of the hydrocarbon mixture. A hydrocarbon includes a kinematic viscosity according to ASTM D7042 and/or ASTM D445 of from about 5 mm²/s to about 8 mm²/s, preferably from about 5 to about 7 mm²/s at a temperature ranging from about −35° C. to about −33° C. Each of these attributes work synergistically to provide for better fuel efficiency while being used by a rocket in a combustion process.

As described above, a disclosed rocket propellant includes a liquid oxygen and a hydrocarbon mixture. A liquid oxygen may be provided by any source such as a cooled oxygen, compressed oxygen, or both so that it is in a substantially liquid state. A hydrocarbon mixture may be derived from a natural source such as from a wellbore and may be synthesized such as through a Fischer-Tropsch reaction. A rocket propellant may have these components in various amounts with respect to each other. For example, a rocket propellant may have a ratio of a liquid oxygen to a hydrocarbon mixture ranging from about 1 to about 5. In some embodiments, a rocket propellant contains a liquid oxygen and a hydrocarbon in a mixture ratio of about 1, or about 1.25, or about 1.5, or about 1.75, or about 2, or about 2.25, or about 2.5, or about 2.75, or about 3, or about 3.25, or about 3.5, or about 3.75, or about 4, where about includes plus or minus 0.125.

As disclosed above, a rocket propellant includes a hydrocarbon. A hydrocarbon may include, but is not limited to, a paraffin, a straight chain paraffin, a branched paraffin, a cyclo paraffin, and combinations thereof. A hydrocarbon may include alkanes, alkenes, alkynes, aromatic, saturated, unsaturated molecules, and combinations thereof. In some embodiments, a rocket propellant composition may include a hydrocarbon mixture having a hydrogen content according to ASTM D5291 ranging from about 14.0 mass % to about 16.0 mass %, preferably from about 15.0 to about 16.0 mass % by mass of the hydrocarbon mixture. Hydrogen content may be due to hydrogen atoms covalently bonded to hydrocarbon contained within a hydrocarbon mixture. In some embodiments, a rocket propellant composition may include a hydrogen content of about 14.0 mass %, or about 14.1 mass %, or about 14.2 mass %, or about 14.3 mass %, or about 14.4 mass %, 14.5 mass %, or about 14.5 mass %, or about 14.6 mass %, or about 14.7 mass %, or about 14.8 mass %, or about 14.9 mass % or about 15.0 mass %, or about 15.1 mass %, or about 15.2 mass %, or about 15.3 mass %, or about 15.4 mass %, or about 15.5 mass %, or about 15.6 mass %, or about 15.7 mass %, or about 15.8 mass %, or about 15.9 mass %, or about 16.0 mass %, by mass of a hydrocarbon mixture, where about includes plus or minus 0.05 mass %.

A hydrocarbon mixture of a disclosed embodiment may have a flash point of greater than about 60° C. at 1 ATM. A hydrocarbon mixture may have a flash point ranging from about 60° C. to about 100° C. at 1 ATM. In some embodiments, a hydrocarbon mixture may have a flash point of about 60° C., or about 65° C., or about 70° C., or about 75° C., or about 80° C., or about 85° C., or about 90° C., or about 95° C., or about 100° C. at 1 ATM, where about includes plus or minus 2.5° C.

A hydrocarbon mixture may have a distillation initial boiling point of at least about 180° C. at 1 ATM. According to some embodiments, a hydrocarbon mixture may have a distillation initial boiling point according to ASTM D86 ranging from about 160° C. to about 200° C. at 1 ATM, preferably ranging from 180° C. to 200° C. at 1 ATM. For example, a hydrocarbon mixture may have an initial boiling of about 180° C., or about 182.5° C., or About 185° C., or about 185.5° C., or about 190° C., or about 192.5° C., or About 195° C., or about 195.5° C., or about 200° C. at 1 ATM, where about includes plus or minus 1.25° C.

A disclosed hydrocarbon mixture may have a density according to ASTM D4052 ranging from about 0.5 kg/m³ to about 1 kg/m³ at a temperature of about 15° C. In some embodiments, a hydrocarbon mixture may have a density of about 0.5 kg/m³, or about 0.6 kg/m³, or about 0.7 kg/m³, or about 0.8 kg/m³, or about 0.9 kg/m³, or about 1 kg/m³, at a temperature of about 15° C., where about includes plus or minus 0.05 kg/m³.

In some embodiments, a hydrocarbon mixture may have a freezing point according to ASTM D5792 that is lower than about −45° C. In some embodiments, a hydrocarbon mixture may have a freezing point from about −45° C. to about −90° C. For example, a hydrocarbon mixture can have a freezing point of about −45° C., or about −50° C., or about −55° C., or about −60° C., or about −65° C., or about −70° C., or about −75° C., or about −80° C., or about −85° C., or about −90° C., where about includes plus or minus 2.5° C.

Besides hydrogen, a hydrocarbon mixture may also include sulfur. In some embodiments, a disclosed hydrocarbon mixture contains a sulfur content ranging from about 0.01 ppm to about 0.1 ppm. A hydrocarbon mixture may have a sulfur content of about 0.01 ppm, or about 0.02 ppm, or about 0.03 ppm, or about 0.04 ppm, or about 0.05 ppm, or about 0.06 ppm, or about 0.07 ppm, or about 0.08 ppm, or about 0.09 ppm, or about 0.1 ppm, by mass of the hydrocarbon mixture, where about includes plus or minus 0.05 ppm.

According to some embodiments, a rocket propellant composition may have a specific impulse value of at least about 1% gain versus RP-1/RP-2 under set operating conditions. In some embodiments, a method of measuring a specific impulse is to measure a characteristic exit velocity (C*). C* is related to specific impulse (I_(sp)) through a standard gravity (g) and constant C_(f), which is dependent on geometries of an engine and heat transfer effects. A formula for specific impulse is I_(SP)=(C*XC_(f))/g. C* can be calculated by using the formula C*=(PXA_(t))/{dot over (m)}. P is a pressure of a combustion chamber, A_(t) is a nozzle throat area, and m is a fuel mass flow rate. In some embodiments, a disclosed rocket propellant composition includes a specific impulse of greater than about 5,700 ft/s at 750 psi and MR=2.7.

Methods of Producing a Rocket Propellant Composition

In some embodiments, methods of producing a rocket propellant composition may include the steps of (a) providing a liquid oxygen and (b) providing a hydrocarbon mixture having a hydrogen content according to ASTM D5291 from about 14.0 mass % to about 16.0 mass %, preferably from about 15 mass % to about 16 mass % by mass of the hydrocarbon mixture. A hydrocarbon mixture may include kinematic viscosity according to ASTM D7042 and/or ASTM D445 of from about 5 mm²/s to about 8 mm²/s, preferably of from about 5 mm²/s to about 7 mm²/s at a temperature ranging from about −35° C. to about −33° C. and a sulfur content of from about 0.01 ppm to about 0.1 mass %, by mass of the hydrocarbon mixture, according to some embodiments. A method may include combining a liquid oxygen and a hydrocarbon mixture and then combusting the combination.

As described above, a method may include combining a liquid oxygen and a hydrocarbon mixture. Combining a liquid oxygen and a hydrocarbon may be done at a burner plate or at some point of combustion. A disclosed method may include combining a liquid oxygen and a hydrocarbon mixture at a ratio ranging from about 1 to about 4. For example, a method may include combining a liquid oxygen and a hydrocarbon mixture at a mixture ratio of about 1, or about 1.25, or about 1.5, or about 1.75, or about 2, or about 2.25, or about 2.5, or about 2.75, or about 3, or about 3.25, or about 3.5, or about 3.75, or about 4, where about includes plus or minus 0.125.

EXAMPLES

The following examples illustrate some specific example embodiments of the present disclosure. These examples represent specific approaches found to function well in the practice of the application, and thus can be considered to constitute examples of modes for its practice. However, those of skill in the art should, in light of the present disclosure, appreciate that many changes can be made in the specific embodiments that are disclosed without departing from the spirit and scope of the application.

Example 1

A plot was made of the percent change in useful payload versus the percent change in the specific impulse (I_(SP)) (FIG. 1 ). The plot was made using the Ideal Rocket Equation applied to a 2-stage SpaceX Falcon 9 launch to low Earth orbit. Standard assumptions were used, such as an increase in the final velocity to account for the lack of air resistance and gravitational effects in the Ideal Rocket Equation. As shown in FIG. 1 , an about 1% I_(sp) increase corresponds to an about 5% increase in useful payload. This figure should be treated as indicative rather than precise; but highlights how sensitive the useful payload is to the specific impulse metric (driven by the chemistry and combustion properties of the propellants identified in this application).

Example 2

One hydrocarbon mixture sample (GTL product GS190) was produced and compared by assessing their specific impulse gain at various liquid oxygen:hydrocarbon mixture ratios (FIGS. 2-4 ). Three plots were made plotting the effective exhaust velocity c* (an experimental measurement of the exhaust gas velocity from a rocket engine, related directly to Isp through the gravitational constant g) versus the liquid oxygen:hydrocarbon mixture ratio at three pressures of 350 psi (FIG. 2 ), 500 psi (FIG. 3 ), and 750 psi (FIG. 4 ). FIG. 3 shows Sample 1 lower- and upper-bounds (from real test data) at 350 psi. FIG. 4 compares Sample 1 against RP-1 lower- and upper-bounds (from real test data) at 500 psi, and FIG. 5 compares RP-2 and Sample 1 at 750 psi. These figures show how the novel rocket fuel disclosed in this application, Sample 1, shows Isp performance improvements (through the c* measurement) versus RP-1/RP-2, these becoming more pronounced at higher combustion chamber pressures and at the higher mixture ratio values shown.

Example 3

Kinematic viscosity data was obtained for four rocket propellant samples, which are Sample 1, RP-1, and RP-2 (FIG. 5 ). This data demonstrates how disclosed rocket propellant sample (Sample 1) has advantageous cryogenic cooling capabilities over known rocket propellants (RP-1, RP-2) as the viscosity for Sample 1 does not reach the standard viscosity of RP-1 or RP-2, 4.5 cSt, until significantly cooler temperatures than do RP-1 or RP-2.

As shown in FIG. 5 , in comparison, known rocket propellants RP-1 and RP-2 hit the limiting viscosity at a temperature of about −7° C.

Example 4

Two dimensional gas chromatography, or GC×GC data was obtained for a disclosed rocket propellant sample (GS190). As shown in Table 1, 41.92% of the rocket propellant is branched chain paraffins having a 12 carbon length with 25% of the rocket propellant being branched chain paraffins having an 11 carbon length. Besides branched paraffins, 24.57% of the rocket propellant includes linear paraffins. Also as shown in Table 1, 25% are linear paraffins, 71% are branched paraffins, 4% are naphthenics, and 1% are di-naphthenics. The important thing to note with these metrics is the lack of aromatic content: aromatic content in RP-1 (and even more-so in RP-2) is required to be low to reduce the propensity of the fuel to break-down under high thermal stress (coking). The fact that the fuel is primarily comprised of paraffins should enhance the resistance to thermal decomposition or coking, making this rocket fuel an attractive option for rocket engines that rely on the pre-combusted rocket fuel for cooling capability.

TABLE 1 Rocket Propellant GC X GC data. C# nP isoP N DiN Mo Ar NmoAr DiAr NdiAr TriAr Total <5 0 0 0 5 0.00 0.00 0.00 0.00 6 0.00 0.00 0.00 0.00 0.00 0.00 7 0.00 0.00 0.00 0.00 0.00 0.00 8 0.00 0.00 0.00 0.00 0.00 0.00 9 0.01 0.00 0.00 0.00 0.01 0.00 0.02 10 1.24 0.25 0.14 0.05 0.00 0.00 0.00 1.68 11 17.73 24.66 2.03 0.26 0.00 0.00 0.00 44.69 12 5.31 41.92 1.24 0.34 0.00 0.00 0.00 0.00 48.82 13 0.20 4.09 0.14 0.00 0.00 0.00 0.00 0.00 4.44 14 0.05 0.20 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.25 15 0.02 0.06 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.08 16 0.01 0.01 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.01 17 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 18 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 19 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 20 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 21 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 22 0.01 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.01 >22 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 0.00 Total 24.57 71.20 3.56 0.66 0.02 0.00 0.00 0.00 0.00 100.00 KEY nP Normal (linear) Paraffins isoP Iso (branched) Paraffins N Naphthenics DiN Di-Naphthenics Mo Ar Mono-Aromatics NmoAr Naphthenic-mono-Aromatics DiAr Di-Aromatics NdiAr Naphthenic-di-Aromatics TriAr Tri-Aromatics

It is understood that the listed components for each unit are for illustration purposes only, and this is not intended to limit the scope of the application. A specific combination of these or other components or units can be configured in such a composition or method for the intended use based on the teachings in the application.

Persons skilled in the art may make various changes in the shape, size, number, separation characteristic, and/or arrangement of parts without departing from the scope of the instant disclosure. Each disclosed component, system, and process step may be performed in association with any other disclosed component, system, or process step and in any order according to some embodiments. Where the verb “may” appears, it is intended to convey an optional and/or permissive condition, but its use is not intended to suggest any lack of operability unless otherwise indicated. Persons skilled in the art may make various changes in methods of preparing and using a composition, device, and/or system of the disclosure. Where desired, some embodiments of the disclosure may be practiced to the exclusion of other embodiments.

Also, where ranges have been provided, the disclosed endpoints may be treated as exact and/or approximations as desired or demanded by the particular embodiment. Where the endpoints are approximate, the degree of flexibility may vary in proportion to the order of magnitude of the range. For example, on one hand, a range endpoint of about 50 in the context of a range of about 5 to about 50 may include 50.5, but not 52.5 or 55 and, on the other hand, a range endpoint of about 50 in the context of a range of about 0.5 to about 50 may include 55, but not 60 or 75. In addition, it may be desirable, in some embodiments, to mix and match range endpoints. Also, in some embodiments, each figure disclosed (e.g., in one or more of the examples, tables, and/or drawings) may form the basis of a range (e.g., depicted value+/−about 10%, depicted value+/−about 50%, depicted value+/−about 100%) and/or a range endpoint. With respect to the former, a value of 50 depicted in an example, table, and/or drawing may form the basis of a range of, for example, about 45 to about 55, about 25 to about 100, and/or about 0 to about 100.

These equivalents and alternatives along with obvious changes and modifications are intended to be included within the scope of the present disclosure. Accordingly, the foregoing disclosure is intended to be illustrative, but not limiting, of the scope of the disclosure as illustrated by the appended claims.

The title, abstract, background, and headings are provided in compliance with regulations and/or for the convenience of the reader. They include no admissions as to the scope and content of prior art and no limitations applicable to all disclosed embodiments. 

What is claimed is:
 1. A rocket propellant composition comprising: (a) a liquid oxygen; and (b) a hydrocarbon mixture comprising: (i) a hydrogen content from about 14.0 mass % to about 16.0 mass %, by mass of the hydrocarbon mixture; (ii) a kinematic viscosity of from about 5 mm²/s to about 8 mm²/s at a temperature ranging from about −35° C. to about −33° C.; and (iii) a sulfur content of from about 0 01 ppm to about 0.1 ppm, by mass of the hydrocarbon mixture.
 2. The rocket propellant composition according to claim 1, wherein a mixture ratio of the liquid oxygen to the hydrocarbon mixture ranges from about 1.8 to about 3.2.
 3. The rocket propellant composition according to claim 1, wherein a mixture ratio of the liquid oxygen to the hydrocarbon mixture ranges from about 1 to about
 4. 4. The rocket propellant composition according to claim 1, wherein the hydrogen content ranges from about 15.3 mass % to about 15.9 mass %, by mass of the hydrocarbon mixture.
 5. The rocket propellant composition according to claim 1, wherein the hydrocarbon mixture further comprises one or more of a branched paraffin, a straight chain paraffin, a cyclo-paraffin, and combinations thereof.
 6. The rocket propellant composition according to claim 1, wherein the hydrocarbon further comprises a flash point according to ASTM D93 of at least about 60° C. at 1 ATM.
 7. The rocket propellant composition according to claim 1, wherein the hydrocarbon mixture further has a distillation initial boiling point of at least about 180° C.
 8. The rocket propellant composition according to claim 1, wherein the hydrocarbon mixture further has a density ranging from about 0.7 kg/m3 to about 0.9 kg/m3 at a temperature of about 15° C.
 9. The rocket propellant composition according to claim 1, wherein the hydrocarbon mixture further comprises a freezing point that is lower than about −45° C.
 10. The rocket propellant composition according to claim 1, wherein the hydrocarbon mixture is a Fischer-Tropsch product.
 11. The rocket propellant composition according to claim 1, wherein the rocket propellant composition comprises a c* value of at least about 1% gain versus RP-1/RP-2 under set operating conditions.
 12. A method for producing a rocket propellant composition, the method comprising: (a) providing a liquid oxygen; and (b) providing a hydrocarbon mixture comprising: (i) a hydrogen content from about 15.0 mass % to about 16.0 mass %, by mass of the hydrocarbon mixture; (ii) a kinematic viscosity of from about 5 mm²/s to about 7 mm²/s at a temperature ranging from about −35° C. to about −33° C.; and (iii) a sulfur content of from about 0.01 ppm to about 0.1 mass %, by mass of the hydrocarbon mixture.
 13. The method according to claim 12, wherein a mixture ratio of the liquid oxygen to the hydrocarbon mixture ranges from about 1 to about
 4. 14. The method according to claim 12, wherein a mixture ratio of the liquid oxygen to the hydrocarbon mixture ranges from about 1.8 to about 3.2.
 15. The method according to claim 12, further comprising combining the liquid oxygen and the hydrocarbon mixture at a burner plate.
 16. The method according to claim 12, wherein the hydrogen content ranges from about 15.3 mass % to about 15.9 mass %, by mass of the hydrocarbon mixture.
 17. The method according to claim 12, wherein the hydrocarbon further comprises a flash point according to ASTM D93 of at least about 60° C. at 1 ATM.
 18. The method according to claim 12, wherein the hydrocarbon mixture further comprises a distillation initial boiling point of at least about 190° C.
 19. The method according to claim 12, wherein the rocket fuel composition further comprises a density ranging from about 0.7 kg/m3 to about 0.8 kg/m3 at a temperature of about 15° C.
 20. The method according to claim 12, wherein the hydrocarbon mixture is a Fischer-Tropsch product.
 21. The method according to claim 12, wherein the rocket propellant composition comprises a c* value of at least about 1% gain versus RP-1/RP-2 under set operating conditions. 